Reinforced composite materials are increasingly used as structural elements in modern aircraft. In particular, panels of composite material are now available for use as external surfaces on aircraft such as wing surfaces and fuselage surfaces. These panels can be manufactured in large sizes and then bonded to underlying composite or non-composite support structures such as stringers, spars, ribs, runners, etc. A typical composite panel includes a plurality of layers of woven cloth having a high tensile strength, such as a cloth made from carbon fibers with intervening layers of a resinous material, such as epoxy. The entire laminate is then cured under elevated temperature and pressure to form a consolidated composite structure. This consolidated structure can be engineered with anisotropic load-bearing properties by varying the number and direction of the fibers, if desired, and is thus highly desirable for high-strength, low-weight applications in aircraft.
Composites of this type are subject to damage, such as ballistic damage, when used on military aircraft. Repairing composite panels is difficult, particularly in the field, because of the nature of composites. Specifically, a composite panel which has been damaged is preferably repaired in such a manner to ensure that the repaired surface is flush with the surrounding, undamaged area of the panel, and to maintain the anisotropic load-bearing properties of the surrounding panel. Such repairs are further complicated by the fact that the panel may form part of a closed surface such as a portion of a wing, in which the interior surface or "blind side" of the panel is inaccessible. Further yet, the repair of such panels generally requires that the damaged area be trimmed so that the resulting aperture and patch have complimentary chamfered or tapered edges between an angle of 1.degree.-5.degree.. Such angle is typically referred to as "scarf angle" or taper angle.
Presently, blind side repairs of composite panels entail the use of grinding or milling steps. Dust and grit may adversely affect bonding of the patch to the trimmed area. Grinding or milling is labor intensive, and often fails to produce a smooth taper or scarf at the desired angle. For example, on a typical composite panel having a thickness on the order of approximately 1/2", the damaged area is milled away using a conventional milling apparatus by milling out concentric steps of increasingly reduced diameter towards one side of the damaged panel. The milling step is typically preceded by a gross grinding step which generally removes the damaged area from the panel. The milling procedure results in a step-like series of concentric rings of reduced diameter. Corresponding steps are then milled into a repair plug for mating with the milled area in the damaged panel. The resulting junction between the repair plug and the composite panel macroscopically resembles a smooth taper, but does not achieve the optimal gradual taper which is preferred for a proper scarf between the repair plug and the panel.
Various apparatus are available for cutting holes in panels. U.S. Pat. No. 2,793,478, to Rohowetz, and U.S. Pat. No. 4,603,678, to Fish, disclose band saw type devices which are capable of performing circular cuts and plunge cuts, respectively, but neither device is capable of trimming a damaged area of a panel without cutting through the perimeter of the panel. Such breaching of the panel perimeter is undesirable for panel repairs. Grinding apparatus such as the device disclosed in U.S. Pat. No. 2,583,396, to Skoog, would produce undesirable dust and grit as described above.
In view of the above, a need exists for a blind side method for trimming out a damaged area of a substantially closed surface without obtaining access to the "blind side" of the surface.
In addition, a need exists for a method for preparing a plug for a prepared damaged area of a composite panel having the desired scarf angle thereon.